The present disclosure relates to a gas turbine engine and, more particularly, to cooling of a combustor wall assembly and method of design.
Gas turbine engines, such as those that power modem commercial and military aircraft, include a fan section to propel the aircraft, compressor section to pressurize a supply of air from the fan section, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases and generate thrust.
The combustor section may have an annular wall assembly having inner and outer shells that support respective inner and outer heat shielding liners. The liners may be comprised of a plurality of floating heat shields or panels that together define an annular combustion chamber. An annular cooling cavity is defined between the respective shells and liners for supplying cooling air to an opposite hot side of the panels through a plurality of strategically placed effusion holes. Impingement holes are located in the shell for supply cooling air from an outer air plenum and into the cavity. The effusion holes are generally orientated to create a protective blanket, or, air film over the hot side of the panels, thereby protecting the panels from the hot combustion gases in the chamber. Cooling pins may be located in the cavity and project outward from the cold side of the liner to further conduct heat out of the liner.
Unfortunately, the placement of impingement holes, effusion holes and cooling pins relative to one another is somewhat random complicating wall assembly design with respect to cooling and leading to less than optimal use of cooling air that may reduce engine efficiency.